US4248566A - Dual function compressor bleed - Google Patents

Dual function compressor bleed Download PDF

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Publication number
US4248566A
US4248566A US05/949,143 US94914378A US4248566A US 4248566 A US4248566 A US 4248566A US 94914378 A US94914378 A US 94914378A US 4248566 A US4248566 A US 4248566A
Authority
US
United States
Prior art keywords
rotor
meridional
shroud
bleed
flow
Prior art date
Legal status (The legal status is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the status listed.)
Expired - Lifetime
Application number
US05/949,143
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English (en)
Inventor
Dennis C. Chapman
David T. Sayre
Current Assignee (The listed assignees may be inaccurate. Google has not performed a legal analysis and makes no representation or warranty as to the accuracy of the list.)
Rolls Royce Corp
JPMorgan Chase Bank NA
Original Assignee
Motors Liquidation Co
Priority date (The priority date is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the date listed.)
Filing date
Publication date
Application filed by Motors Liquidation Co filed Critical Motors Liquidation Co
Priority to US05/949,143 priority Critical patent/US4248566A/en
Priority to CA328,555A priority patent/CA1111008A/en
Priority to DE19792935235 priority patent/DE2935235A1/de
Priority to IT50277/79A priority patent/IT1164902B/it
Priority to GB7932632A priority patent/GB2032523B/en
Priority to JP12694579A priority patent/JPS5551990A/ja
Priority to FR7924954A priority patent/FR2438181A1/fr
Application granted granted Critical
Publication of US4248566A publication Critical patent/US4248566A/en
Assigned to AEC ACQUISTION CORPORATION reassignment AEC ACQUISTION CORPORATION ASSIGNMENT OF ASSIGNORS INTEREST (SEE DOCUMENT FOR DETAILS). Assignors: GENERAL MOTORS CORPORATION
Assigned to CHEMICAL BANK, AS AGENT reassignment CHEMICAL BANK, AS AGENT ASSIGNMENT OF ASSIGNORS INTEREST (SEE DOCUMENT FOR DETAILS). Assignors: AEC ACQUISITION CORPORATION
Assigned to ALLISON ENGINE COMPANY, INC. reassignment ALLISON ENGINE COMPANY, INC. CHANGE OF NAME (SEE DOCUMENT FOR DETAILS). Assignors: AEC ACQUISTITION CORPORATION A/K/A AEC ACQUISTION CORPORATION
Anticipated expiration legal-status Critical
Expired - Lifetime legal-status Critical Current

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Classifications

    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F04—POSITIVE - DISPLACEMENT MACHINES FOR LIQUIDS; PUMPS FOR LIQUIDS OR ELASTIC FLUIDS
    • F04D—NON-POSITIVE-DISPLACEMENT PUMPS
    • F04D27/00—Control, e.g. regulation, of pumps, pumping installations or pumping systems specially adapted for elastic fluids
    • F04D27/02—Surge control
    • F04D27/0207—Surge control by bleeding, bypassing or recycling fluids
    • F04D27/0215—Arrangements therefor, e.g. bleed or by-pass valves
    • Y—GENERAL TAGGING OF NEW TECHNOLOGICAL DEVELOPMENTS; GENERAL TAGGING OF CROSS-SECTIONAL TECHNOLOGIES SPANNING OVER SEVERAL SECTIONS OF THE IPC; TECHNICAL SUBJECTS COVERED BY FORMER USPC CROSS-REFERENCE ART COLLECTIONS [XRACs] AND DIGESTS
    • Y10—TECHNICAL SUBJECTS COVERED BY FORMER USPC
    • Y10S—TECHNICAL SUBJECTS COVERED BY FORMER USPC CROSS-REFERENCE ART COLLECTIONS [XRACs] AND DIGESTS
    • Y10S415/00—Rotary kinetic fluid motors or pumps
    • Y10S415/914—Device to control boundary layer

Definitions

  • This invention relates to centrifugal and mixed flow compressors for use in gas turbine engines and more particularly to centrifugal compressors including gas bleed in association therewith for regulating the operating characteristics of the compressor.
  • Type 2 waveoffs are more specifically ones where there is a snap deceleration from full power in the engine followed immediately by a snap acceleration from near idle speed of the engine.
  • inadequate surge margin in such compressors could be eliminated by bleeding a substantial percentage of the compressor gas flow through a bleed valve connected to communicate with the compressor discharge scroll.
  • bleed valves can open in the normal operating range of the gas turbine engine and cause the engine to be energy inefficient.
  • an object of the present invention is to improve the operating efficiency of compressors in gas turbine engines by including dual purpose bleed means therein to move the inducer stall to lower speeds and lower flow conditions in the compressor and moreover to enhance full speed flow capabilities of the compressor.
  • Another object of the present invention is to provide improved operating efficiency by reduction of inducer stall to lower speed and lower flow conditions while maintaining enhanced full speed flow capabilities through the compressor by including a bleed control slot at a meridional point of the stationary cover over the impeller downstream of the inducer choke point where the compressor impeller produces an in-flow of gas through the slot at compressor speeds near the compressor impeller design speed thereby to add the flow through the control slot to that of the annular inlet area to the impeller to increase total in-flow of gas to the impeller under high speed conditions of operation; and wherein the control slot is operative to bleed gas from the compressor exteriorly thereof at speeds less than design speed of the impeller to flow stabilize the impeller at part speed phases of its operation.
  • Still another object of the present invention is to provide an improved flow controlled centrifugal compressor for use in gas turbine engines wherein the compressor includes a stationary shroud cover with a control slot located circumferentially therearound at the impeller tip at a particular meridional length aft of the inducer leading edge of the compressor and wherein the slot is sized and located aft of the flow limiting throat restriction in the inducer of the impeller so that under inducer choking conditions the inducer flow limiting restriction of flow is supplemented by in-flow of gas through the control slot to in-bleed sufficient gas through gas bleed aft of the throat restriction to add to the high speed flow capacity of the compressor and wherein the same control slot serves to produce an adequate out-flow of inlet gas flow through the inducer at part speed conditions of operation of the impeller to improve the part speed surge characteristics of the compressor.
  • FIG. 1 is a fragmentary, longitudinal sectional view partially in elevation of a compressor including the present invention
  • FIG. 2 is a fragmentary, sectional view through a portion of a stationary shroud cover in FIG. 1;
  • FIG. 3 is a fragmentary, enlarged elevational view of the shroud in FIG. 2;
  • FIG. 4 is a front elevational view of the shroud cover of the present invention.
  • FIG. 5 is a performance chart of a compressor with and without the present invention.
  • FIG. 6 is a chart of cover static pressure versus corrected speed of operation of the impeller in the compressor of FIG. 1.
  • FIG. 1 a compressor 10 is shown. It includes a front support assembly 12 and a rear support assembly 14 for physically locating the rotary components of the compressor 10 in a manner to be discussed. More particularly, the front support assembly 12 includes a plurality of circumferentially spaced axial struts 16 located in a generally radial direction across an annular inlet 18 to a rotor assembly 20 interposed between the front support assembly 12 and the rear support assembly 14.
  • the front support assembly 12 includes an outer annular shroud wall 22 having a stepped shoulder 24 on the downstream end thereof that is piloted with respect to the forward flange 26 of a stationary compressor outer shroud housing 28.
  • the front assembly shroud wall 22 has a contour that defines a smooth pathed outer surface 30 of an axially extending inlet flow path 32 that prevents abrupt flow changes upstream of a contoured inner surface 34 of the stationary compressor shroud housing 28.
  • the front support assembly 12 includes an internal hub portion 36 of conoidal form that defines a smooth transition to the inlet 18 and further defines a smooth contoured inner annular wall 38 that likewise avoids abrupt flow changes through the inlet flow path 32 to the contoured hub surface 40 on an impeller hub 42 of the rotor assembly 20.
  • the airflow path through the assembly is thereby arranged to produce as uniform a flow distribution as possible from the inlet 18 to a flow inducer core 44 of the rotor assembly 20.
  • the flow inducer core 44 is made up of a plurality of full length rotor impeller blades 46 having inducer passages 47 therebetween.
  • a plurality of flow splitter blades 48 are also included on hub 42.
  • Each of the full blades 46 includes a leading edge 50.
  • the full blades 46 each have a radially outwardly located contoured tip 52 that follows the contour of the inner surface of a liner 54 of abradable aluminum compound that minimizes the operating clearance between the rotor assembly 20 and the inner surface of the stationary shroud housing 28.
  • each of the full blades is bent back tangentially from the radial direction at an outlet radius 56 of the rotor assembly 20.
  • Each of the splitter blades 48 includes a leading edge 58 and a contoured radially outer tip 60 that is shaped to coincide with a contour of the liner 54.
  • Each of the splitter blades is likewise bent back tangentially from the radial direction at the outlet radius 60 of the rotor assembly 20.
  • the rotor assembly 20 is fixed for rotation with respect to the inner contoured surface of the liner 54 by a rear bearing assembly 64 and a front bearing assembly 66.
  • the rear bearing assembly 64 supportingly receives a rear hub extension 68 having a bore therethrough that receives a splined adapter 72 having internal spline teeth 74 thereon and an end portion 76 fixedly secured to the hub 42 by a suitable fastener represented by the illustrated screw and lock washer combination 78.
  • a compressor drive shaft can be coupled to the splined adapter 72 for driving the rotor assembly 20 during compressor operation.
  • the rear bearing assembly 64 further includes a roller bearing 80 supported in a bearing support 82 of the rear support assembly 14.
  • the support assembly 14 includes a pair of axially spaced abradable seal lands 84, 86 that cooperate with labyrinth seals 88, 90 on the impeller hub 42 to seal the internal gas flow path through the compressor assembly 10 from low pressure cavities within the compressor.
  • the rear support assembly 14 includes an internal surface 92 forming the back of the compressor rearwardly of the rear wall 94 of the impeller hub 42 as shown in FIG. 1.
  • a pilot flange 96 on assembly 14 supports the rear wall 98 of a diffuser 100 configured to receive discharge flow from rotor assembly 20.
  • the diffuser 100 includes a front wall 102 that is secured to a pilot flange 104 on the outer radius of the stationary compressor shroud housing 28 by a plurality of fasteners 106 located at circumferentially spaced points around the shroud. Fasteners 106 further secure a discharge scroll collector 108 to the outlet 110 of the diffuser 100.
  • the diffuser 100 includes a leading inlet edge 112 that is spaced to the outer radius of the rotor assembly 20 as shown in FIG. 1 to define an inlet region 114 in which shock wave patterns can develop to require matching the flow patterns through the rotor assembly 20 and those through the diffuser assembly 100.
  • the most crucial part of any high mach number diffuser is the inlet region or space 114. This quasivaneless region is complicated by the presence of shock waves therein and substantial sidewall boundary layer build up.
  • the diffuser 100 is correlated with respect to the high performance impeller characteristics including choke-to-surge operating ranges of the compressor 10, pressure recovery through the diffuser 100, and total pressure loss. Parameters that affect these variables include the number of diffuser passages, diffuser area ratio, diffuser passage length, leading edge to impeller tip radius ratio and the diffuser entry region geometry among others. While vaned diffusers are shown in the illustrated embodiment, the invention is equally applicable to any diffuser construction.
  • compressor pressure ratio is plotted as a function of flow rate through the compressor with lines of constant rotational speed being superimposed thereon.
  • the line connecting the left hand terminus of each of the illustrated speed lines is called the surge line and represents a limit of aerodynamically stable operation in centrifugal compressor and diffuser assemblies.
  • the higher flow end of the surge line 116 is typically determined by stall within the diffuser of a compressor in a gas turbine engine, such as diffuser 100 illustrated in FIG. 1.
  • the lower flow end of the surge line is typically determined by a combination of diffuser stall and stall within the inducer or inlet end of the rotor assembly 20.
  • the low speed region in which the inducer leading edge 50 reaches stall conditions is frequently characterized by a dip in the surge line, shown at 115 in FIG. 5.
  • a dip in the surge line may occur at a sufficiently high value of flow and speed of operation of the compressor to preclude energy efficient correction by means of a compressor discharge bleed valve system.
  • Another way to relieve such problems where compressor discharge bleed is an unacceptable compromise of performance in a primary engine operating region is to relieve inducer stall by the location of an air bleed through holes or slots in the stationary cover over the inducer.
  • FIG. 6 is a plot of static pressure measured at various meridional distances along the inner contoured surface 34 of the stationary compressor shroud housing 28 from the inlet to the exit thereof.
  • the indicated static pressure measurements have been corrected to standard inlet conditions and are plotted as a function of compressor speed along an engine operating line.
  • a horizontal line 118 on the chart of FIG. 6 represents the pressure level of 14.7 psi, in this case considered to be an ambient condition.
  • the preselected rotor assembly 20 and vaned diffuser 100, in the present invention are characterized as producing a given static pressure profile inside the cover at a meridional length 119 along the cover 28 with a zero percent point 121 at the leading edge of inducer 44 and its 100 percent point 124 at the exit edge 56.
  • Low meridional distances are those near the point 121 just inside the shroud.
  • Intermediate meridional distances are in the range of 10 to 15% of length 119.
  • Higher meridional distances are in the range of 30% and above of length 119.
  • the static pressure levels at the 15 percent meridional distance along the contoured surface 34 are both above and below 14.7 psia.
  • a continous slot 120 is located at the 15 percent meridional point on shroud 28 to produce an airflow which flows outwardly from within the inducer passages 47 to the exterior of the compressor 10 for all speeds up to approximately 95 percent of the design speed. Above 95 percent of the design speed of the rotor assembly, however, the pressure differential is actually reversed from outside of the compressor 10 to the inducer passages 47 at the contoured wall 34 thereof so that flow enters the inducer through a continuous bleed slot 120.
  • inducers having the configuration of that illustrated in the present invention normally tend to limit flow capacity of a centrifugal compressor at and above 100 percent of its design speed with a vaned diffuser and at all speeds with conventional vaneless diffusers.
  • a throat or flow limiting restriction in the inducer is upstream of the illustrated 15 percent meridional distance location of slot 120.
  • the potential of inbleed air through the bleed slot 120 downstream of the restriction throat in such an inducer of centrifugal compressors constitutes a way of adding high speed flow capacity to high performance centrifugal compressors.
  • the high speed inbleed capability through the slot 120 further, may be used to decrease the annulus size of the annular inlet flow path 32 which, under part speed operating conditions, further improves the part speed stall and surge characteristics of such compressors.
  • the slot 120 is formed completely around the circumference of the shroud cover 28 at the previously discussed 15 percent meridional point.
  • the slot is spanned by raised bridges 122 on the shroud cover 28 which serve to carry structural loads. If desired, separate holes in a circumferential row could replace slot 120 as long as desired bleed flow area is maintained.
  • the improved dual function control slot 120 or equivalent holes constitute a static device which, by virtue of its strategic location, produces variable flow patterns in a compressor for a gas turbine engine to extend its surge range and to improve its operating efficiency.
  • the variability of bleed direction and improved results therefrom eliminates the need for control valves that close to prevent power reducing air loss when the surge control bleed is not required.

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  • Engineering & Computer Science (AREA)
  • Life Sciences & Earth Sciences (AREA)
  • Sustainable Development (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Structures Of Non-Positive Displacement Pumps (AREA)
  • Control Of Positive-Displacement Air Blowers (AREA)
US05/949,143 1978-10-06 1978-10-06 Dual function compressor bleed Expired - Lifetime US4248566A (en)

Priority Applications (7)

Application Number Priority Date Filing Date Title
US05/949,143 US4248566A (en) 1978-10-06 1978-10-06 Dual function compressor bleed
CA328,555A CA1111008A (en) 1978-10-06 1979-05-29 Dual function compressor bleed
DE19792935235 DE2935235A1 (de) 1978-10-06 1979-08-30 Gasverdichter
IT50277/79A IT1164902B (it) 1978-10-06 1979-09-17 Compressore a flusso centrifugo e misto per motori a turbina a gas
GB7932632A GB2032523B (en) 1978-10-06 1979-09-20 Controlled flow gas compressor
JP12694579A JPS5551990A (en) 1978-10-06 1979-10-03 Controlled flow gas compressor
FR7924954A FR2438181A1 (fr) 1978-10-06 1979-10-08 Compresseur de gaz a debit commande pour moteur a turbine a gaz

Applications Claiming Priority (1)

Application Number Priority Date Filing Date Title
US05/949,143 US4248566A (en) 1978-10-06 1978-10-06 Dual function compressor bleed

Publications (1)

Publication Number Publication Date
US4248566A true US4248566A (en) 1981-02-03

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US05/949,143 Expired - Lifetime US4248566A (en) 1978-10-06 1978-10-06 Dual function compressor bleed

Country Status (7)

Country Link
US (1) US4248566A (de)
JP (1) JPS5551990A (de)
CA (1) CA1111008A (de)
DE (1) DE2935235A1 (de)
FR (1) FR2438181A1 (de)
GB (1) GB2032523B (de)
IT (1) IT1164902B (de)

Cited By (43)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
EP0092955A2 (de) * 1982-04-22 1983-11-02 A/S Kongsberg Väpenfabrikk Verfahren und Vorrichtung zur Beeinflussung der Grenzschicht des Mediums in einem Kompressor
EP0229519A1 (de) * 1985-12-24 1987-07-22 Holset Engineering Company Limited Kompressoren
US4687412A (en) * 1985-07-03 1987-08-18 Pratt & Whitney Canada Inc. Impeller shroud
US4930978A (en) * 1988-07-01 1990-06-05 Household Manufacturing, Inc. Compressor stage with multiple vented inducer shroud
US4930979A (en) * 1985-12-24 1990-06-05 Cummins Engine Company, Inc. Compressors
US5236301A (en) * 1991-12-23 1993-08-17 Allied-Signal Inc. Centrifugal compressor
US5235803A (en) * 1992-03-27 1993-08-17 Sundstrand Corporation Auxiliary power unit for use in an aircraft
US5277541A (en) * 1991-12-23 1994-01-11 Allied-Signal Inc. Vaned shroud for centrifugal compressor
US5295785A (en) * 1992-12-23 1994-03-22 Caterpillar Inc. Turbocharger having reduced noise emissions
US5601406A (en) * 1994-12-21 1997-02-11 Alliedsignal Inc. Centrifugal compressor hub containment assembly
WO2000046509A1 (en) 1999-02-04 2000-08-10 Pratt & Whitney Canada Corp. Compressor bleeding using an uninterrupted annular slot
US6325595B1 (en) 2000-03-24 2001-12-04 General Electric Company High recovery multi-use bleed
US6699008B2 (en) 2001-06-15 2004-03-02 Concepts Eti, Inc. Flow stabilizing device
US20050152775A1 (en) * 2004-01-14 2005-07-14 Concepts Eti, Inc. Secondary flow control system
US20050163606A1 (en) * 2004-01-22 2005-07-28 Svihla Gary R. Centrifugal compressor with channel ring defined inlet recirculation channel
US20050232762A1 (en) * 2004-04-20 2005-10-20 Honeywell International Inc. Turbomachine compressor scroll with load-carrying inlet vanes
US7025356B1 (en) * 2004-12-20 2006-04-11 Pratt & Whitney Canada Corp. Air-oil seal
US20070110568A1 (en) * 2005-11-17 2007-05-17 Honeywell International, Inc. Pilot relief to reduce strut effects at pilot interface
US20080069690A1 (en) * 2006-09-18 2008-03-20 Pratt & Whitney Canada Corp. Thermal and external load isolating impeller shroud
US20090232642A1 (en) * 2008-03-12 2009-09-17 Atte Anema Adjustable compressor bleed system and method
US20100028148A1 (en) * 2007-06-06 2010-02-04 Akihiro Nakaniwa Sealing device for rotary fluid machine, and rotary fluid machine
US20100111688A1 (en) * 2008-10-30 2010-05-06 Honeywell International Inc. Axial-centrifugal compressor with ported shroud
US20100119367A1 (en) * 2007-06-06 2010-05-13 Akihiro Nakaniwa Sealing device for rotary fluid machine, and rotary fluid machine
US20110129332A1 (en) * 2008-07-01 2011-06-02 Snecma Axial-centrifugal compressor having system for controlling play
US20120102969A1 (en) * 2010-10-28 2012-05-03 Wagner Joel H Centrifugal compressor with bleed flow splitter for a gas turbine engine
WO2012078443A1 (en) * 2010-12-08 2012-06-14 Alcoa Inc. Locking nut assambly
US20140202202A1 (en) * 2012-03-22 2014-07-24 Panasonic Corporation Centrifugal compressor
US20160177965A1 (en) * 2014-12-17 2016-06-23 Electro-Motive Diesel, Inc. Compressor assembly for turbocharger burst containment
US20160195099A1 (en) * 2013-07-18 2016-07-07 Snecma Cover of a turbomachine centrifugal compressor capable of being rigidly connected via the downstream side near to the upstream edge of same, and turbomachine comprising this cover
US20160281727A1 (en) * 2015-03-27 2016-09-29 Dresser-Rand Company Apparatus, system, and method for compressing a process fluid
US9650916B2 (en) 2014-04-09 2017-05-16 Honeywell International Inc. Turbomachine cooling systems
US9683488B2 (en) 2013-03-01 2017-06-20 Rolls-Royce North American Technologies, Inc. Gas turbine engine impeller system for an intermediate pressure (IP) compressor
US20180045214A1 (en) * 2016-08-15 2018-02-15 Borgwarner, Inc. Compressor wheel, method of making the same, and turbocharger including the same
US20180087647A1 (en) * 2016-09-23 2018-03-29 Bell Helicopter Textron Inc. Fan with labyrinth seal for prevention of water damage to a gearbox
US10359051B2 (en) 2016-01-26 2019-07-23 Honeywell International Inc. Impeller shroud supports having mid-impeller bleed flow passages and gas turbine engines including the same
US11156226B2 (en) 2018-02-09 2021-10-26 Carrier Corporation Centrifugal compressor with recirculation passage
US11199195B2 (en) * 2019-10-18 2021-12-14 Pratt & Whitney Canada Corp. Shroud with continuous slot and angled bridges
US20220018290A1 (en) * 2020-07-16 2022-01-20 Raytheon Technologies Corporation Gas turbine engine including seal assembly with abradable coating and cutter
US20220018289A1 (en) * 2020-07-16 2022-01-20 Raytheon Technologies Corporation Gas turbine engine including seal assembly with abradable coating including magnetic particles embedded in polymer
US20220018291A1 (en) * 2020-07-16 2022-01-20 Raytheon Technologies Corporation Gas turbine engine including seal assembly with abradable coating including magnetic particles
US20220195888A1 (en) * 2020-12-18 2022-06-23 Pratt & Whitney Canada Corp. Bearing housing assembly
US11680582B2 (en) * 2017-09-25 2023-06-20 Johnson Controls Tyco IP Holdings LLP Two piece split scroll for centrifugal compressor
US11846249B1 (en) * 2022-09-02 2023-12-19 Rtx Corporation Gas turbine engine with integral bypass duct

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FR2904037B1 (fr) * 2006-07-19 2010-11-12 Snecma Ventilation d'une cavite aval de rouet de compresseur centrifuge

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US3123285A (en) * 1964-03-03 Diffuser with boundary layer control
BE509180A (de) * 1951-02-15
US2837270A (en) * 1952-07-24 1958-06-03 Gen Motors Corp Axial flow compressor
US3217655A (en) * 1962-09-04 1965-11-16 Snecma Centrifugal pump
US3484039A (en) * 1967-07-14 1969-12-16 Georg S Mittelstaedt Fans and compressors
US3887295A (en) * 1973-12-03 1975-06-03 Gen Motors Corp Compressor inlet control ring
US3893787A (en) * 1974-03-14 1975-07-08 United Aircraft Corp Centrifugal compressor boundary layer control
SU591619A1 (ru) * 1976-04-08 1978-02-05 Московское Ордена Ленина И Ордена Трудового Красного Знамени Высшее Техническое Училище Им.Н.Э.Баумана Рабочее колесо центробежного компрессора

Cited By (76)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
EP0092955A2 (de) * 1982-04-22 1983-11-02 A/S Kongsberg Väpenfabrikk Verfahren und Vorrichtung zur Beeinflussung der Grenzschicht des Mediums in einem Kompressor
US4479755A (en) * 1982-04-22 1984-10-30 A/S Kongsberg Vapenfabrikk Compressor boundary layer bleeding system
EP0092955A3 (de) * 1982-04-22 1985-12-18 A/S Kongsberg Väpenfabrikk Verfahren und Vorrichtung zur Beeinflussung der Grenzschicht des Mediums in einem Kompressor
US4687412A (en) * 1985-07-03 1987-08-18 Pratt & Whitney Canada Inc. Impeller shroud
EP0229519A1 (de) * 1985-12-24 1987-07-22 Holset Engineering Company Limited Kompressoren
US4743161A (en) * 1985-12-24 1988-05-10 Holset Engineering Company Limited Compressors
US4930979A (en) * 1985-12-24 1990-06-05 Cummins Engine Company, Inc. Compressors
US4930978A (en) * 1988-07-01 1990-06-05 Household Manufacturing, Inc. Compressor stage with multiple vented inducer shroud
US5277541A (en) * 1991-12-23 1994-01-11 Allied-Signal Inc. Vaned shroud for centrifugal compressor
US5236301A (en) * 1991-12-23 1993-08-17 Allied-Signal Inc. Centrifugal compressor
US5235803A (en) * 1992-03-27 1993-08-17 Sundstrand Corporation Auxiliary power unit for use in an aircraft
US5295785A (en) * 1992-12-23 1994-03-22 Caterpillar Inc. Turbocharger having reduced noise emissions
US5399064A (en) * 1992-12-23 1995-03-21 Caterpillar Inc. Turbocharger having reduced noise emissions
US5601406A (en) * 1994-12-21 1997-02-11 Alliedsignal Inc. Centrifugal compressor hub containment assembly
US5613830A (en) * 1994-12-21 1997-03-25 Alliedsignal Inc. Centrifugal compressor hub containment assembly
WO2000046509A1 (en) 1999-02-04 2000-08-10 Pratt & Whitney Canada Corp. Compressor bleeding using an uninterrupted annular slot
US6183195B1 (en) 1999-02-04 2001-02-06 Pratt & Whitney Canada Corp. Single slot impeller bleed
US6325595B1 (en) 2000-03-24 2001-12-04 General Electric Company High recovery multi-use bleed
US6699008B2 (en) 2001-06-15 2004-03-02 Concepts Eti, Inc. Flow stabilizing device
US7025557B2 (en) 2004-01-14 2006-04-11 Concepts Eti, Inc. Secondary flow control system
US20050152775A1 (en) * 2004-01-14 2005-07-14 Concepts Eti, Inc. Secondary flow control system
US20050163606A1 (en) * 2004-01-22 2005-07-28 Svihla Gary R. Centrifugal compressor with channel ring defined inlet recirculation channel
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Also Published As

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JPS5551990A (en) 1980-04-16
GB2032523B (en) 1983-03-02
DE2935235C2 (de) 1987-07-30
FR2438181B1 (de) 1984-04-20
GB2032523A (en) 1980-05-08
IT1164902B (it) 1987-04-15
IT7950277A0 (it) 1979-09-17
FR2438181A1 (fr) 1980-04-30
CA1111008A (en) 1981-10-20
DE2935235A1 (de) 1980-04-17

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